A gas turbine engine is a type of internal combustion engine having an upstream rotating compressor coupled to a downstream turbine, with a combustion chamber positioned between the compressor and the turbine.
Incoming air flows through the compressor that brings it to higher pressure. Energy is then added by spraying fuel into the compressed air and igniting it, with the resulting combustion generating a high temperature, high pressure gas flow. This gas flows enters the turbine, where it expands down to the exhaust pressure, while producing a shaft work output in the process. The turbine shaft work is used to drive the compressor and any ancillary devices that may be coupled to the shaft. The remaining energy in the exhaust gases is exhausted from the engine and provides thrust.
Each of the compressor and turbine comprises one or more rotors, with each rotor being paired with a corresponding stator vane array.
In the event of an engine failure that results in a loss of torque carrying capability in one of the rotors, the turbine will experience a transient overspeed event. This will result in a significantly increased mechanical loading of the turbine disc and may cause the turbine disc to burst.
Conventional design protocols require the turbine disc to be sized to withstand such a transient overspeed event. This increases the weight of the turbine disc.
Current aerospace regulatory frameworks stipulate that turbine disc failure events must occur at a rate less than 1×10−8 events per engine flying hour. Current design protocols result in turbine disc designs in which any failure cannot result in the assembly reaching a sufficiently high overspeed as to burst the disc. Such turbine disc designs result in increased turbine disc weight in order to increase its margin of safety.
It is desirable therefore to provide a means of limiting the turbine disc speed reached during an overspeed event, in order to be able to reduce turbine disc weight and so improve its reliability.